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| 1 | Numerical study of corner separation in a linear compressor cascade using various turbulence models显示文摘Three-dimensional corner separation is a common phenomenon that significantly affects compressor performance.Turbulence model is still a weakness for RANS method on predicting corner separation flow accurately.In the present study,numerical study of corner separation in a linear highly loaded prescribed velocity distribution(PVD) compressor cascade has been investigated using seven frequently used turbulence models.The seven turbulence models include Spalart–Allmaras model,standard k–e model,realizable k–e model,standard k–x model,shear stress transport k–x model,v^2–f model and Reynolds stress model.The results of these turbulence models have been compared and analyzed in detail with available experimental data.It is found the standard k–e model,realizable k–e model,v^2–f model and Reynolds stress model can provide reasonable results for predicting three dimensional corner separation in the compressor cascade.The Spalart–Allmaras model,standard k–x model and shear stress transport k–x model overestimate corner separation region at incidence of 0°.The turbulence characteristics are discussed and turbulence anisotropy is observed to be stronger in the corner separating region. | Liu Yangwei Yan Hao Liu Yingjie Lu Lipeng Li Qiushi | 2016 | Chinese Journal of Aeronautics2016，29，3： | 11 |
| 2 | Large-eddy simulation of shock-wave/turbulent boundary layer interaction with and without Spark Jet control显示文摘The efficiency and mechanism of an active control device ‘‘Spark Jet' and its application in shock-induced separation control are studied using large-eddy simulation in this paper.The base flow is the interaction of an oblique shock-wave generated by 8° wedge and a spatially-developing Ma = 2.3 turbulent boundary layer.The Reynolds number based on the incoming flow property and the boundary layer displacement thickness at the impinging point without shock-wave is20000.The detailed numerical approaches were presented.The inflow turbulence was generated using the digital filter method to avoid artificial temporal or streamwise periodicity.The numerical results including velocity profile,Reynolds stress profile,skin friction,and wall pressure were systematically validated against the available wind tunnel particle image velocimetry(PIV) measurements of the same flow condition.Further study on the control of flow separation due to the strong shock-viscous interaction using an active control actuator ‘‘Spark Jet' was conducted.The single-pulsed characteristic of the device was obtained and compared with the experiment.Both instantaneous and time-averaged flow fields have shown that the jet flow issuing from the actuator cavity enhances the flow mixing inside the boundary layer,making the boundary layer more resistant to flow separation.Skin friction coefficient distribution shows that the separation bubble length is reduced by about 35% with control exerted. | Yang Guang Yao Yufeng Fang Jian Gan Tian Li Qiushi Lu Lipeng | 2016 | Chinese Journal of Aeronautics2016，29，3： | 8 |
| 3 | Modified k-ω model using kinematic vorticity for corner separation in compressor cascades显示文摘A new method of modifying the conventional k-ω turbulence model for corner separation is proposed in this paper. The production term in the ω equation is modified using kinematic vorticity considering fluid rotation and deformation in complex geometric boundary conditions. The corner separation flow in linear compressor cascades is calculated using the original k-ω model, the modified k-ω model and the Reynolds stress model(RSM). The numerical results of the modified model are compared with the available experimental data, as well as the corresponding results of the original k-ω model and RSM. In terms of accuracy, the modified model, which significantly improves the performance of the original k-ω model for predicting corner separation, is quite competitive with the RSM. However, the modified model, which has considerably lower computational cost, is more robust than the RSM. | LIU YangWei YAN Hao FANG Le LULiPeng LI QiuShi SHAO Liang | 2016 | Science China(Technological Sciences)2016，59，5： | 7 |
| 4 | 引入压力梯度的针对角区分离流动的Spalart-Allmaras模型的改进显示文摘针对Spalart-Allmaras(S-A)模型在角区分离计算中的问题,将无量纲的压力梯度引入其涡黏性输运方程的生成项,得到了改进的S-A模型.通过对两套含角区分离的低速压气机叶栅进行验证计算发现:与实验结果相比,原始S-A模型所得的分离区偏大,分离区内壁面压力偏低;而改进模型得到了与实验一致的分离区尺寸以及吸力面、压力面压力系数分布等结果.针对S-A模型涡黏性生成项和耗散项的分析表明:引入的无量纲压力梯度有效的识别了角区分离,在分离区内改变了涡黏性的生成、耗散关系,增大了涡黏性,从而缩小了计算所得分离区,同时在主流区保留了原始S-A模型的计算结果,进而带来了良好的改进效果. | 马力 孙槿静 陆利蓬 | 2016 | 航空动力学报2016，31，10： | 4 |
| 5 | Numerical simulation of compression corner flows at Mach number 9显示文摘A hypersonic vehicle encounters a wide range of conditions during its complete flight regime.These flight conditions may vary from low to high Mach numbers with varying angles of attack.The near-wall viscous dissipation associated with flows at combined high Mach and Reynolds numbers leads to significant wall heat transfer rates and shear stresses.The shock wave/boundary-layer interaction results in a flow separation region,which commonly augments total pressure losses in the flow and lowers the efficiency of aerodynamic control surfaces such as fins installed on a vehicle.The standard turbulence models,when used to resolve such flows,result in incorrect separation bubble size for large separated flows.Therefore,it results in an inaccurate aerodynamic load,such as the wall pressures,skin friction distribution,and heat transfer rate.In previous studies,the application of the shock-unsteadiness correction to the standard two-equation k-ωturbulence model improved the separation bubble size leading to an accurate pressure prediction and shock definition with the assumption of constant Prandtl number.In the present work,the new shock-unsteadiness modification to the k-ωturbulence model is applied to the hypersonic compression corner flows.This new model with variable Prandtl number is based on the model parameter,which depends upon the local density ratio.The computed wall pressures,heat flux and flow field are compared to the experimental data.A parametric study is carried out by varying compression deflection angles,free stream Reynolds number and wall temperatures to compute the flow field and wall data accurately,particularly in the shock boundary layer interaction region.The new shockunsteadiness modified k-ωmodel with variable Prandtl number shows an accurate prediction of initial pressure rise location,pressure distribution in the plateau region and heat flux in comparison to the standard k-ωmodel. | Amjad APASHA Khalid AJUHANY | 2020 | Chinese Journal of Aeronautics2020，33，6： | 2 |
| 6 | 部分处理机匣的性能预估模型显示文摘利用进气畸变平行压气机理论划分子压气机的思想,将部分处理机匣压气机沿周向分为处理机匣子压气机和实壁机匣子压气机,再由动量守恒和能量守恒求得部分处理机匣压气机出口参数,并依据失速点流量与处理区周向角度呈线性关系的实验结论提出部分处理机匣性能预估模型假设.模型预估与实验对比结果表明:预估特性与实验特性曲线在整体趋势上一致.跨声速条件下98%设计转速峰值效率预估偏差小于0.5%,综合裕度预估偏差小于0.6%. | 朱振坤 袁巍 李秋实 李志平 | 2017 | 航空动力学报2017，32，5： | 1 |
| 7 | Drag increment induced by a small-scale forward-facing step in Mach number5 turbulent boundary layer flows显示文摘Small-scale roughness elements or imperfections are inevitable over the surface of a flight vehicle.The aerodynamics of these small-scale structures is difficult to predict but may play an important role in the design of a flight vehicle at high speed.The forward-facing step is a typical type of roughness element.Many experiments have been conducted to study the aerodynamics of supersonic forward-facing step,especially with a step height larger than boundary layer thickness.However,few studies focus on small steps.To improve the understanding of small-scale forwardfacing step flow,we perform a series of simulations to analyze its aerodynamic influence on a Mach number 5 turbulent boundary layer.The general flow structures are analyzed and discussed.Several shock waves can be induced by the step even if the step height is much smaller than the boundary layer thickness.Two significant shocks are the separation shock and the reattachment shock.The influenced area by the step is limited.With the increase of the step height,the non-dimensional influence area decreases and gradually converges when the step height reaches the boundary layer thickness.There are two normalized distributions of the skin friction coefficient and pressure coefficient associated with step height.By using the normalized parameters,a power-law relationship between the step height and the drag increment coefficient is revealed and fits the simulation results well.It is further illustrated that this relationship still holds when changing the inlet angle of attack,but needs slight modification with the angle of attack. | Yifei XUE Jie REN Jinling LUO Song FU | 2020 | Chinese Journal of Aeronautics2020，33，10： | 1 |
| 8 | 民用飞机风扇的计算气动弹性综述(英文)显示文摘The aim of this paper is to present the state-of-the art in computational aeroelasticity methods that are available for analyzing fan blades on modern civil aircraft. Fan blades in modern high-bypass aero-engines typically produce around 80% of the thrust. In order to improve specific fuel consumption and reduce the level of noise emitted from the engine, civil turbofan engine designs are moving toward even larger fan diameters with lower tip speeds and hence the importance of this component of aero-engine becomes even more prominent. To reduce weight, future fan blades will be made of composite materials and shorter intakes are used. The new designs are highly loaded and will be more susceptible to aerodynamic and aeroelastic instabilities, and hence computationally efficient aeroelastic modelling tools for such blades are paramount. | M.Vahdati K.B Lee S.Stapelfeldt | 2018 | 风机技术2018，60，5： | 0 |
| 9 | 机载布撒亚音速子弹药初始外弹道性能研究显示文摘为获得亚音速下不同弹体外形对机载布撒子弹药气动特性和初始外弹道性能的影响规律,采用FLUENT流体动力学计算软件对亚音速流中不同弹体外形的子弹药气动性能进行了仿真分析,建立了机载布撒子弹药的6自由度刚体外弹道分析模型,并基于龙格-库塔法求解其初始外弹道性能。研究结果表明:弹体外形影响下气体绕流表现出的层流/湍流行为是影响弹体所受气动力的重要因素。弹体头部突起对弹体气动特性影响不大;在0.72 Ma来流中,圆头弹型阻力系数和静力矩系数导数分别为0.349和0.86,相应为平头弹型的31.4%和46.7%;在总翼展面积相同时,增加尾翼数量会减小弹体所受气动升力,8片尾翼弹型的静力矩系数为6片尾翼弹形的58%;初始外弹道飞行时凹头弹型和6片尾翼弹型的攻角收敛更快,径向转动惯量增大时攻角收敛速率减慢。研究结果可为布撒武器子弹药的相关设计提供一定参考。 | 王尧 毛亮 何筱 魏晨杨 祁宇轩 姜春兰 | 2023 | 兵器装备工程学报2023，44，3： | 0 |
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